Turbomachines and epicyclic gear assemblies with symmetrical compound arrangement

ABSTRACT

A gear assembly for use with a turbomachine comprises a sun gear, a plurality of planet gear layshafts that each support a first stage planet gear and a second stage planet gear, and a ring gear. The sun gear is configured to rotate about a longitudinal centerline of the gear assembly, and the plurality of planet gear layshafts have a tubular shape.

ACKNOWLEDGMENT OF GOVERNMENT SUPPORT

The project leading to this application has received funding from theClean Sky 2 Joint Undertaking (JU) under grant agreement No 945541. TheJU receives support from the European Union's Horizon 2020 research andinnovation programme and the Clean Sky 2 JU members other than theUnion.

CROSS REFERENCE TO RELATED APPLICATIONS

This application claims the benefit of Italian Application No.102020000014206, filed Jun. 15, 2020 and Italian Application No.102020000016168, filed Jul. 3, 2020, both of which are incorporatedherein by reference.

BACKGROUND

A turbofan engine operates on the principle that a central gas turbinecore drives a bypass fan, the bypass fan being located at a radiallocation between a nacelle of the engine and the engine core. With sucha configuration, the engine is generally limited in a permissible sizeof the bypass fan, as increasing a size of the fan correspondinglyincreases a size and weight of the nacelle.

An open rotor engine, by contrast, operates on the principle of havingthe bypass fan located outside of the engine nacelle. This permits theuse of larger rotor blades able to act upon a larger volume of air thana traditional turbofan engine, potentially improving propulsiveefficiency over conventional turbofan engine designs.

Engine designs for turbomachines, including turbofans and especiallyopen rotor engines, may require large gear ratios between the low speedspool and the fan rotor to permit the larger rotor blades to act upon alarger volume of air and/or to do so at certain desired operating speedsof the engine or aircraft. One challenge is that known gear assembliesmay provide inadequate gear ratios for desired operations. For example,known gear assemblies may inadequately reduce the output rotationalspeed relative to the input rotational speed, such that the fan rotoroperates too fast and inefficient and/or the turbine operates too slowand inefficient.

As such, there is a need for gear assemblies that provide desired gearratios as may be suitable for turbomachines, including open rotorengines.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the technology disclosed in the description.

Various turbomachine engines and gear assemblies are disclosed herein.In some embodiments, a turbomachine engine that includes a fan assemblyand a core engine comprising a turbine and an input shaft rotatable withthe turbine is provided. A single-stage epicyclic gear assembly receivesthe input shaft at a first speed and drives an output shaft coupled tothe fan assembly at a second speed, the second speed being slower thanthe first speed. The gear assembly comprises a sun gear, a plurality ofplanet gears, and a ring gear. The sun gear rotates about a longitudinalcenterline of the gear assembly and has a sun gear-mesh region along thelongitudinal centerline of the gear assembly where the sun gear isconfigured to contact the plurality of planet gears. A ring gear-meshregion is provided along the longitudinal centerline of the gearassembly where the ring gear is configured to contact the plurality ofplanet gears. The sun gear-mesh region is axially offset from the ringgear-mesh region along the longitudinal centerline.

These and other features, aspects and advantages of the presentdisclosure will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the disclosed technology and, together with thedescription, serve to explain the principles of the disclosure.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention directed to oneof ordinary skill in the art, is set forth in the specification, whichmakes reference to the appended figures, in which:

FIG. 1 is a cross-sectional schematic illustration of an exemplaryembodiment of an open rotor propulsion system;

FIG. 2 is a cross-sectional schematic illustration of an exemplaryembodiment of a ducted propulsion system;

FIG. 3 is a schematic illustration of an exemplary gear assembly;

FIG. 4 is a cross-sectional view of an exemplary planet gear layshaft;

FIG. 5 is a schematic illustration of an exemplary planet gear layshaftwith first and second stage planet gears;

FIG. 6 is a schematic illustration of an exemplary gear assembly withcompound symmetry;

FIG. 7 is a schematic illustration of an exemplary gear assembly withcompound symmetry, with a portion of a ring gear removed for clarity;

FIG. 8 is a schematic illustration of an exemplary planet carrier with acompound planet gear.

FIG. 9 is another schematic illustration of an exemplary planet carrierwith a compound planet gear;

FIG. 10 is an aft-side view of an exemplary gear assembly.

FIG. 11 is a cross-sectional view of a portion of an exemplary gearassembly.

DETAILED DESCRIPTION

Reference now will be made in detail to embodiments of the invention,one or more examples of which are illustrated in the drawings. Eachexample is provided by way of explanation of the invention, notlimitation of the invention. In fact, it will be apparent to thoseskilled in the art that various modifications and variations can be madein the present invention without departing from the scope or spirit ofthe invention. For instance, features illustrated or described as partof one embodiment can be used with another embodiment to yield a stillfurther embodiment. Thus, it is intended that the present inventioncovers such modifications and variations as come within the scope of theappended claims and their equivalents.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

Referring now to the drawings, FIG. 1 is an exemplary embodiment of anengine 100 including a gear assembly 102 according to aspects of thepresent disclosure. The engine 100 includes a fan assembly 104 driven bya core engine 106. In various embodiments, the core engine 106 is aBrayton cycle system configured to drive the fan assembly 104. The coreengine 106 is shrouded, at least in part, by an outer casing 114. Thefan assembly 104 includes a plurality of fan blades 108. A vane assembly110 is extended from the outer casing 114. The vane assembly 110including a plurality of vanes 112 is positioned in operable arrangementwith the fan blades 108 to provide thrust, control thrust vector, abateor re-direct undesired acoustic noise, and/or otherwise desirably altera flow of air relative to the fan blades 108. In some embodiments, thefan assembly 104 includes between three (3) and twenty (20) fan blades108. In particular embodiments, the fan assembly 104 includes betweenten (10) and sixteen (16) fan blades 108. In certain embodiments, thefan assembly 104 includes twelve (12) fan blades 108. In certainembodiments, the vane assembly 110 includes an equal or fewer quantityof vanes 112 to fan blades 108.

In some embodiments, the fan blade tip speed at a cruise flightcondition can be 650 to 900 fps, or 700 to 800 fps. A fan pressure ratio(FPR) for the fan assembly 104 can be 1.04 to 1.10, or in someembodiments 1.05 to 1.08, as measured across the fan blades at a cruiseflight condition.

In certain embodiments, such as depicted in FIG. 1, the vane assembly110 is positioned downstream or aft of the fan assembly 104. However, itshould be appreciated that in some embodiments, the vane assembly 110may be positioned upstream or forward of the fan assembly 104. In stillvarious embodiments, the engine 100 may include a first vane assemblypositioned forward of the fan assembly 104 and a second vane assemblypositioned aft of the fan assembly 104. The fan assembly 104 may beconfigured to desirably adjust pitch at one or more fan blades 108, suchas to control thrust vector, abate or re-direct noise, and/or alterthrust output. The vane assembly 110 may be configured to desirablyadjust pitch at one or more vanes 112, such as to control thrust vector,abate or re-direct noise, and/or alter thrust output. Pitch controlmechanisms at one or both of the fan assembly 104 or the vane assembly110 may co-operate to produce one or more desired effects describedabove.

The core engine 106 is generally encased in outer casing 114 defining amaximum diameter. In certain embodiments, the engine 100 includes alength from a longitudinally forward end 116 to a longitudinally aft end118. In various embodiments, the engine 100 defines a ratio of length(L) to maximum diameter (D_(max)) that provides for reduced installeddrag. In one embodiment, L/D_(max) is at least 2. In another embodiment,L/D_(max) is at least 2.5. In some embodiments, the L/D_(max) is lessthan 5, less than 4, and less than 3. In various embodiments, it shouldbe appreciated that the L/D_(max), is for a single unducted rotorengine.

The reduced installed drag may further provide for improved efficiency,such as improved specific fuel consumption. Additionally, oralternatively, the reduced drag may provide for cruise altitude engineand aircraft operation at or above Mach 0.5. In certain embodiments, theL/D_(max), the fan assembly 104, and/or the vane assembly 110 separatelyor together configure, at least in part, the engine 100 to operate at amaximum cruise altitude operating speed between approximately Mach 0.55and approximately Mach 0.85.

Referring again to FIG. 1, the core engine 106 extends in a radialdirection R relative to an engine axis centerline 120. The gear assembly102 receives power or torque from the core engine 106 through a powerinput source 122 and provides power or torque to drive the fan assembly104, in a circumferential direction C about the engine axis centerline120, through a power output source 124.

In certain embodiments, such as depicted in FIG. 1, the engine 100 is anun-ducted thrust producing system, such that the plurality of fan blades108 is unshrouded by a nacelle or fan casing. As such, in variousembodiments, the engine 100 may be configured as an unshrouded turbofanengine, an open rotor engine, or a propfan engine. In particularembodiments, the engine 100 is a single unducted rotor engine includinga single row of fan blades 108. The engine 100 configured as an openrotor engine includes the fan assembly 104 having large-diameter fanblades 108, such as may be suitable for high bypass ratios, high cruisespeeds (e.g., comparable to aircraft with turbofan engines, or generallyhigher cruise speed than aircraft with turboprop engines), high cruisealtitude (e.g., comparable to aircraft with turbofan engines, orgenerally higher cruise speed than aircraft with turboprop engines),and/or relatively low rotational speeds. Cruise altitude is generally analtitude at which an aircraft levels after climb and prior to descendingto an approach flight phase. In various embodiments, the engine isapplied to a vehicle with a cruise altitude up to approximately 65,000ft. In certain embodiments, cruise altitude is between approximately28,000 ft. and approximately 45,000 ft. In still certain embodiments,cruise altitude is expressed in flight levels (FL) based on a standardair pressure at sea level, in which a cruise flight condition is betweenFL280 and FL650. In another embodiment, cruise flight condition isbetween FL280 and FL450. In still certain embodiments, cruise altitudeis defined based at least on a barometric pressure, in which cruisealtitude is between approximately 4.85 psia and approximately 0.82 psiabased on a sea level pressure of approximately 14.70 psia and sea leveltemperature at approximately 59 degrees Fahrenheit. In anotherembodiment, cruise altitude is between approximately 4.85 psia andapproximately 2.14 psia. It should be appreciated that in certainembodiments, the ranges of cruise altitude defined by pressure may beadjusted based on a different reference sea level pressure and/or sealevel temperature.

Although depicted above as an unshrouded or open rotor engine in FIG. 1,it should be appreciated that the gear assemblies disclosed herein maybe applied to shrouded or ducted engines, partially ducted engines,aft-fan engines, or other turbomachine configurations, including thosefor marine, industrial, or aero-propulsion systems. In addition, thegear assemblies disclosed herein may also be applicable to turbofan,turboprop, or turboshaft engines.

For example, FIG. 2 is a cross-sectional schematic illustration of anexemplary embodiment of an engine 200 that includes the gear assembly202 in combination with a ducted fan propulsion system. However, unlikethe open rotor configuration of FIG. 2, the fan assembly 204 and its fanblades 208 are contained within an annular fan case 230 and the vaneassembly 210 and the vanes 212 extend radially between the fan cowl 232and the inner surface of the fan case 230. As discussed above, the gearassemblies disclosed herein can provide for increased gear ratios for afixed gear envelope (e.g., with the same size ring gear), oralternatively, a smaller diameter ring gear may be used to achieve thesame gear ratios.

As shown in FIG. 2, the core engine 206 is generally encased in outercasing 214, and has a length extending from a longitudinally forward end216 to a longitudinally aft end 218. The exemplary core engine (for aducted or unducted engine) can include a compressor section 240, a heataddition system 242 (e.g., combustor), and an expansion section 244together in serial flow arrangement. The core engine 206 extendscircumferentially relative to an engine centerline axis 220. The coreengine 206 includes a high-speed spool that includes a high-speedcompressor and a high-speed turbine operably rotatably coupled togetherby a high-speed shaft. The heat addition system 232 is positionedbetween the high-speed compressor and the high-speed turbine. Variousembodiments of the heat addition system 232 include a combustionsection. The combustion section may be configured as a deflagrativecombustion section, a rotating detonation combustion section, a pulsedetonation combustion section, or other appropriate heat additionsystem. The heat addition system 232 may be configured as one or more ofa rich-burn system or a lean-burn system, or combinations thereof. Instill various embodiments, the heat addition system 232 includes anannular combustor, a can combustor, a cannular combustor, a trappedvortex combustor (TVC), or other appropriate combustion system, orcombinations thereof.

The core engine 206 can also include a booster or low-speed compressorpositioned in flow relationship with the high-speed compressor. Thelow-speed compressor is rotatably coupled with the low-speed turbine viaa low-speed shaft 246 to enable the low-speed turbine to drive thelow-speed compressor. The low-speed shaft 246 is also operably connectedto gear assembly 202 to provide power to the fan assembly 204 via thepower input source 222, such as described further herein.

It should be appreciated that the terms “low” and “high”, or theirrespective comparative degrees (e.g., -er, where applicable), when usedwith compressor, turbine, shaft, or spool components, each refer torelative speeds within an engine unless otherwise specified. Forexample, a “low turbine” or “low-speed turbine” defines a componentconfigured to operate at a rotational speed, such as a maximum allowablerotational speed, lower than a “high turbine” or “high-speed turbine” atthe engine. Alternatively, unless otherwise specified, theaforementioned terms may be understood in their superlative degree. Forexample, a “low turbine” or “low-speed turbine” may refer to the lowestmaximum rotational speed turbine within a turbine section, a “lowcompressor” or “low speed compressor” may refer to the lowest maximumrotational speed turbine within a compressor section, a “high turbine”or “high-speed turbine” may refer to the highest maximum rotationalspeed turbine within the turbine section, and a “high compressor” or“high-speed compressor” may refer to the highest maximum rotationalspeed compressor within the compressor section. Similarly, the low-speedspool refers to a lower maximum rotational speed than the high-speedspool. It should further be appreciated that the terms “low” or “high”in such aforementioned regards may additionally, or alternatively, beunderstood as relative to minimum allowable speeds, or minimum ormaximum allowable speeds relative to normal, desired, steady state, etc.operation of the engine.

As discussed in more detail below, the core engine includes the gearassembly that is configured to transfer power from the expansion sectionand reduce an output rotational speed at the fan assembly relative to alow-speed turbine. Embodiments of the gear assemblies depicted anddescribed herein can allow for gear ratios suitable for large-diameterunducted fans (e.g., FIG. 1) or certain turbofans (e.g., FIG. 2).Additionally, embodiments of the gear assemblies provided herein may besuitable within the radial or diametrical constraints of the core enginewithin the outer casing.

The gear assemblies described herein includes a gear set for decreasingthe rotational speed of the fan assembly relative to the low speed(pressure) turbine. In operation, the rotating fan blades are driven bythe low speed (pressure) turbine via gear assembly such that the fanblades rotate around the engine axis centerline and generate thrust topropel the engine, and hence an aircraft on which it is mounted, in theforward direction.

Referring again to FIG. 1, it may be desirable that either or both ofthe fan blades 104 or the vanes 112 incorporate a pitch change mechanismsuch that the blades can be rotated with respect to an axis of pitchrotation either independently or in conjunction with one another. Suchpitch change can be utilized to vary thrust and/or swirl effects undervarious operating conditions, including to provide a thrust reversingfeature which may be useful in certain operating conditions such as uponlanding an aircraft.

Vanes 112 can be sized, shaped, and configured to impart a counteractingswirl to the fluid so that in a downstream direction aft of both fanblades 104 and vanes 112 the fluid has a greatly reduced degree ofswirl, which translates to an increased level of induced efficiency.Vanes 112 may have a shorter span than fan blades 104. For example, insome embodiments, vanes 112 may have a span that is at least 50% of aspan of fan blades 104. In other embodiments, the span of the vanes canbe the same or longer than the span as fan blades 104, if desired. Vanes112 may be attached to an aircraft structure associated with the engine100, as shown in FIG. 1, or another aircraft structure such as a wing,pylon, or fuselage. Vanes 112 may be fewer or greater in number than, orthe same in number as, the number of fan blades 104. In someembodiments, the number of vanes 112 are greater than two, or greaterthan four, in number. Fan blades 104 may be sized, shaped, and contouredwith the desired blade loading in mind.

FIGS. 1 and 2 illustrate what may be termed a “puller” configurationwhere the fan assembly is located forward of the engine core. Otherconfigurations are possible and contemplated as within the scope of thepresent disclosure, such as what may be termed a “pusher” configurationembodiment where the engine core is located forward of the fan assembly.The selection of “puller” or “pusher” configurations may be made inconcert with the selection of mounting orientations with respect to theairframe of the intended aircraft application, and some may bestructurally or operationally advantageous depending upon whether themounting location and orientation are wing-mounted, fuselage-mounted, ortail-mounted configurations.

In the exemplary embodiment of FIG. 1, air enters the core engine 106through an opening aft of the fan blades. FIG. 2 illustrates such anopening in more detail. As shown in FIG. 2, air A entering the fanassembly is divided between air that enters the core engine A1 and airthat bypasses the core engine A2.

Embodiments of the gear assemblies depicted and described herein mayprovide for gear ratios and arrangements that fit within the L/D_(max)constraints of the engine. In certain embodiments, the gear assembliesdepicted and described allow for gear ratios and arrangements providingfor rotational speed of the fan assembly corresponding to one or moreranges of cruise altitude and/or cruise speed provided above.

Various embodiments of the gear assemblies provided herein can allow forgear ratios of up to 14:1. Still various embodiments of the gearassembly can allow for gear ratios of at least 6:1. Still yet variousembodiments of the gear assembly provided herein allow for gear ratiosbetween 6:1 to 12:1, between 7:1 and 11:1, and between 8:1 and 10:1. Itshould be appreciated that embodiments of the gear assembly providedherein may allow for large gear ratios and within constraints such as,but not limited to, length (L) of the engine 10, maximum diameter(D_(max)) of the engine, cruise altitude of up to 65,000 ft, and/oroperating cruise speed of up to Mach 0.85, or combinations thereof.

Various exemplary gear assemblies are shown and described herein. Thesegear assemblies may be utilized with any of the exemplary engines and/orany other suitable engine for which such gear assemblies may bedesirable. In such a manner, it will be appreciated that the gearassemblies disclosed herein may generally be operable with an enginehaving a rotating element with a plurality of rotor blades and aturbomachine having a turbine and a shaft rotatable with the turbine.With such an engine, the rotating element (e.g., fan assembly) may bedriven by the shaft (e.g., low-speed shaft) of the turbomachine throughthe gear assembly.

FIG. 3 illustrates an exemplary gear assembly 302 with a compoundsymmetrical arrangement. Gear assembly 302 includes a sun gear 304 witha diameter D_(s), a plurality of first stage planet gears 306 with adiameter D_(p1), a plurality of second stage planet gears 308 with adiameter D_(p2), and a ring gear 310 with a diameter D_(r). Each of thesun gear 304, planet gears 306, 308, and ring gear 310 are doublehelical gears with first and second sets of helical teeth that areinclined at an acute angle relative to each other. In particular, sungear 204 comprises a first sun gear set 312 and a second sun gear set314. Each of the first stage planet gears comprises a first planet gearset 316 and second planet gear set 318, and each of the second stageplanet gears comprises a third planet gear set 320 and a fourth planetgear set 322. The ring gear 310 comprises a first ring gear set 324 anda second ring gear set 326.

As discussed in more detail below, the number of planet gears can vary.In one embodiment, there are three compound planet gears (e.g., 306,308). In another embodiment, there can be two to four compound planetgears.

As shown in FIG. 4, the compound planet gears 306, 308 are supported bya layshaft 330 that has a tubular configuration. As used herein“tubular” means a longitudinally extending structure that has a hollowinterior section. The tubular layshaft can comprise an inner surface andan outer surface that supports the first and second stage planet gearsas discussed herein.

The tubular layshaft can comprise an intermediate portion 332 thatsupports the first stage planet gears 306 between two outer portions334. As shown in FIG. 4 (and elsewhere), a diameter of the first stageplanet gears D_(p1) can be greater than a diameter of the second stageplanet gears D_(p2). In some embodiments, a ratio of D_(p1): D_(p2) canvary from 1.0 to 2.0, or in some embodiments, from 1.2 to 1.7, or from1.3 to 1.6, or from 1.4 to 1.5.

Because of the difference in diameters of the first and second stageplanet gears, the tubular layshaft 330 can similarly vary to supportthese gears. Thus, as shown in FIG. 4, the intermediate portion 332 hasa greater diameter than the outer portions 334, with angled portions ofthe layshaft extending between the intermediate portion 332 and outerportions 334.

In some embodiments, layshaft 330 can comprise a plurality of holes 338to scavenge lubricating oil within the gear assembly. As shown in FIG.4, the plurality of holes 338 can extend circumferentially around thelayshaft 330 at the intermediate portion 332 (e.g., the enlargedportion) between first and second outer portions 334.

FIG. 5 illustrates the layshaft 330 and compound planet gears 306, 308.FIGS. 6 and 7 illustrate the layshaft 330 and compound planet gears 306,308, with the ring gear 310 (FIG. 6) and with a portion of the ring gearremoved (FIG. 7). In the embodiments shown in FIGS. 6 and 7, threecompound planet gears are provided (306, 308) and the ring gear 310comprises two halves with an interconnecting flanged portion 336. FIG. 7also discloses a plurality of radial passages for oil scavenging.

FIGS. 8 and 9 illustrate a planet carrier 328 with a single compoundplanet gear (306, 308) provided therein for clarity. The layshaft 330extends through an opening in the fore and aft sides of the planetcarrier 328. In some embodiments, the carrier can be connected to theengine frame via a flexible support system, with the flexible supportsystem being configured to collect oil and scavenge oil via holes at alower portion.

FIG. 10 is an aft-side view of gear assembly 302, indicating a pluralityof apertures 360 in the planet carrier 328 to provide assembly access tobolts associated with a flange portion 344 of input shaft 342. Referringto FIGS. 3, 10, and 11, in some embodiments, a bolting diameter of theflange portion of the input shaft can be higher than an innermost arc ofa bore of the layshaft 330. In another embodiment, the bolting diameter346 of the flange portion of the input shaft can be lower than theinnermost arc of the bore of the layshaft 330.

As discussed above, the sun gear can be connected to a turbine via aninput shaft 342. An input shaft can comprise a flexible connectingportion. In some embodiments, the flexible connecting portion caninclude at least a pair of thin flanges bolted together. In someembodiments, the flexible connecting portion can comprise radialpassages for oil scavenge. In some embodiments, the input shaft 342 canbe connected via a spline to the turbine shaft.

In some embodiments, the gear ratio split between the first and secondstages can range from 40% to 60% for each stage (i.e., from 40% to 60%for the first stage and from 60% to 40% for the second stage).

As discussed above, in some embodiments, the sun gear 304, planet gears306, 308, and ring gear 310 can be double helical gears with first andsecond sets of helical teeth that are inclined at an acute anglerelative to each other.

In the embodiment shown in FIG. 3, the gear assembly 302 is a star gearconfiguration in which the planet carrier is generally fixed (e.g.,static) within the engine by support structure. The sun gear 304 isdriven by an input shaft (e.g., a low-speed shaft). A planet gearcarrier 328 is rotatably coupled to a layshaft of the compound planetgears 306, 308, and the ring gear 310 is configured to rotate about thelongitudinal engine axis centerline in a circumferential direction,which in turn drives the power output source (e.g., a fan shaft) that iscoupled to and configured to rotate with the ring gear to drive the fanassembly. In this embodiment, the low-speed shaft rotates in acircumferential direction that is the opposite of the direction in whichthe fan shaft rotates.

In other embodiments, the gear assembly can have a planetaryconfiguration in which the ring gear is fixed (e.g., static) within theengine by a support structure. The sun gear is driven by an input shaft(i.e., low-speed shaft) and instead of the ring gear rotating, theplanet carrier rotates in the same direction of the low-speed shaftrotation direction, to drive the power output source (e.g., a fan shaft)and fan assembly.

Referring again to FIG. 3, the ring gear 310 is coupled to a fan driveshaft 340 to drive the fans. The sun gear 304 is coupled to an inputpower source (e.g., input shaft 342). In some embodiments, the inputshaft can be integrally formed with the sun gear. The bi-helical meshesof the planet gears axially balance the load over the four (phased) gearsets of each compound planet gear. The second stage of planet gears 308can be supported by two rows of cylindrical roller bearings 344 at theplanet bore. In addition, the fan drive shaft 340 can be supported bytapered roller bearings 350, which supports the fan drive shaft in anaxially compact manner. In some embodiments, the roller bearings can beformed from a ceramic material. In some embodiments, the roller bearingscan be lubrication by under-race lubrication, in which lubrication isdirected under the inner race and forced out through a plurality ofholes in an inner race. In some embodiments, as shown in FIG. 3, aninner supporting element of both sets of the roller bearings 344 can bea solid unique elements.

In some embodiments, an oil transfer bearing for a pitch controlmechanism can be provided on a fore side of the gear assembly, and gearassembly/pitch control oil feed lines can be provided through the staticplanet carrier.

In some embodiments, one of the pair of gear sets (e.g., one of thefirst and second gear sets, one of the third or fourth gear sets) isangularly clocked by a set amount of gear pitch relative to the othergear set. For example, the teeth of the first gear set can be angularlyclocked by a first amount of the gear pitch relative to the teeth of thesecond gear set. The first amount can be between one fourth and onehalf. Similarly, the teeth of the third gear set can be angularlyclocked by a second amount of the gear pitch relative to the teeth ofthe fourth gear set. The second amount can be between one fourth and onehalf.

The gear assemblies disclosed herein can provide significant advantagesover conventional systems. For example, the tubular planet gearlayshafts can provide reduced weight gear assemblies, while alsoproviding improve torque reaction in the midplane section of the gearassembly. In addition, relative to conventional systems, assembly ofgearboxes having tubular planet gear layshafts can be more efficient.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The patentable scope ofthe invention is defined by the claims, and may include other examplesthat occur to those skilled in the art. Such other examples are intendedto be within the scope of the claims if they include structural elementsthat do not differ from the literal language of the claims, or if theyinclude equivalent structural elements with insubstantial differencesfrom the literal languages of the claims.

Further aspects of the invention are provided by the subject matter ofthe following clauses:

1. A turbomachine engine (100, 200) comprising a fan assembly (104, 204)comprising a plurality of fan blades (108, 208); and a core engine (106,206) comprising a turbine and an input shaft (122, 222) rotatable withthe turbine; and a gear assembly (102, 202, 302) that receives the inputshaft (122, 222) at a first speed and drives an output shaft (124, 224)coupled to the fan assembly (104, 204) at a second speed, the secondspeed being slower than the first speed. The gear assembly (102, 202,302) comprises a sun gear (304), a plurality of planet gear layshafts(330) that each support a first stage planet gear (306) and a secondstage planet gear (308), and a ring gear (310). The sun gear rotatesabout a longitudinal centerline (120, 220, 320) of the gear assembly andthe plurality of planet gear layshafts have a tubular shape.

2. The turbomachine engine of any of the clauses herein, wherein theplanet gear layshaft (330) comprises a first outer portion (334), asecond outer portion (334), and an intermediate portion (332) betweenfirst and second outer portions, and the first stage planet gear (306)comprises a first gear set (316) and a second gear set (318) supportedat the intermediate portion of the planet gear layshaft, and the secondstage planet gear (308) comprises a third gear set (320) supported atthe first outer portion of the planet gear layshaft and a fourth gearset (322) supported at the second outer portion of the planet gearlayshaft.

3. The turbomachine engine of any of the clauses herein, wherein the sungear (304), the first stage and second stage planet gears (306, 308),and the ring gear (310) comprise double helical gears.

4. The turbomachine engine of any of the clauses herein, wherein thethird and fourth gear set (320, 322) are spaced apart by the first andsecond gear set (316, 318) and the third and fourth gear set (320, 322)are inclined at an acute angle relative to each other.

5. The turbomachine engine of any of the clauses herein, wherein thering gear (310) comprises a first ring gear set (324) that meshes withthe third gear set (320) and a second ring gear set (326) that mesheswith the fourth gear set (322).

6. The turbomachine of any of the clauses herein, wherein a gear ratioof the gear assembly ranges from 6:1 to 14:1, from 6.1 to 12:1, from 7:1to 11:1, or from 8:1 to 10:1.

7. The turbomachine engine of any of the clauses herein, wherein thegear assembly is a star gear configuration in which a planet carriergears is fixed relative to the turbomachine engine and does not rotate.

8. The turbomachine engine of any of the clauses herein, wherein thegear assembly is a planetary gear configuration in which the ring gearis fixed relative to the turbomachine engine and does not rotate.

9. The turbomachine engine of any of the clauses herein, wherein the fanassembly is a single stage of unducted fan blades.

10. The turbomachine engine of any of the clauses herein, wherein thefirst stage planet gear has a first diameter and the second stage planetgear has a second diameter, wherein a ratio of the first diameter to thesecond diameter ranges from 1.0 to 2.0, 1.2 to 1.7, 1.3 to 1.6, or 1.4to 1.5.

11. The turbomachine engine of any of the clauses herein, furthercomprising two rows of roller bearings that support, respectively, thethird gear set and the fourth gear set of the second stage planet gear.

12. The turbomachine engine of any of the clauses herein, wherein thereare three planet gear layshafts (330).

13. A gear assembly (102, 202, 302) for use with a turbomachine, thegear assembly comprising a sun gear (304); a plurality of planet gearlayshafts (330) that each support a first stage planet gear (306) and asecond stage planet gear (308), and a ring gear (310), the planet gearlayshaft (330) comprising a first outer portion (334), a second outerportion (334), and an intermediate portion (332) between first andsecond outer portions. The first stage planet gear (306) comprises afirst gear set (316) and a second gear set (318) supported at theintermediate portion of the planet gear layshaft, and the second stageplanet gear (308) comprises a third gear set (320) supported at thefirst outer portion of the planet gear layshaft and a fourth gear set(322) supported at the second outer portion of the planet gear layshaft,and the plurality of planet gear layshafts have a tubular shape.

14. The turbomachine engine of any of the clauses herein, wherein theteeth of the first gear set are angularly clocked by a first amount of agear pitch with regards to the teeth of the second gear set.

15. The turbomachine engine of clause 14, wherein the first amount isbetween one fourth and one half.

16. The turbomachine engine of any of the clauses herein, wherein teethof the third gear set are angularly clocked by a second amount of thegear pitch with regards to the teeth of the fourth gear set.

17. The turbomachine engine of clause 16, wherein the second amount isbetween one fourth and one half.

18. The turbomachine engine of any of the clauses herein, where therollers are made by ceramic material.

19. The turbomachine engine of any of the clauses herein, wherein therollers are lubricated under race.

20. The turbomachine engine of any of the clauses herein, wherein theinner supporting element of the first and second roller bearings is asolid unique element.

21. The turbomachine engine any of the clauses herein, wherein eachlayshaft has a plurality of scavenge holes in its largest section of aninner tubular cavity.

22. The turbomachine engine of any of the clauses herein, wherein thesun gear is connected to a turbine via an input shaft and wherein atleast one bending flexible element is present.

23. The turbomachine engine of any of the clauses herein, wherein theflexibility of the input shaft is provided, at least in part, via atleast a pair of thin flanges bolted together.

24. The turbomachine engine of any of the clauses herein, wherein theflexible element comprises radial passages for oil scavenge.

25. The turbomachine engine of any of the clauses herein, wherein theinput shaft is connected via a spline to the turbine shaft.

26. The turbomachine engine of any of the clauses herein, wherein thecarrier is connected to the engine frame via a flexible support system.

27. The turbomachine engine of any of the clauses herein, wherein theflexible support system connecting the carrier to the engine frame isconfigured to collect oil and scavenge oil via holes at a lower portion.

28. The turbomachine engine of any of the clauses herein, wherein thebolting diameter of at least one flange in the input shaft is higherthan the innermost arc of the bore of the layshaft pin.

29. The turbomachine engine of any of the clauses herein, wherein thebolting diameter of at least one flange in the input shaft is lower thanthe outermost arc of the bore of the layshaft pin.

30. The turbomachine engine of any of the clauses herein, furthercomprising a plurality of radial passages for oil scavenging in theflanged portion of the ring gear.

We claim:
 1. A turbomachine engine comprising: a fan assembly comprisinga plurality of fan blades; a core engine comprising a turbine and aninput shaft rotatable with the turbine; and a gear assembly thatreceives the input shaft at a first speed and drives an output shaftcoupled to the fan assembly at a second speed, the second speed beingslower than the first speed, the gear assembly comprising a sun gear, aplurality of planet gear layshafts that each support a first stageplanet gear and a second stage planet gear, and a ring gear, the sungear rotating about a longitudinal centerline of the gear assembly,wherein the plurality of planet gear lay shafts have a tubular shape. 2.The turbomachine engine of claim 1, wherein each of the planet gearlayshaft comprises a first outer portion, a second outer portion, and anintermediate portion between first and second outer portions, andwherein the first stage planet gear comprises a first gear set and asecond gear set supported at the intermediate portion of the planet gearlayshaft, and the second stage planet gear comprises a third gear setsupported at the first outer portion of the planet gear layshaft and afourth gear set supported at the second outer portion of the planet gearlayshaft.
 3. The turbomachine engine of claim 2, wherein the sun gear,the first stage and second stage planet gears, and the ring gearcomprise double helical gears.
 4. The turbomachine engine of claim 2,wherein the third and fourth gear set are spaced apart by the first andsecond gear set and the third and fourth gear set are inclined at anacute angle relative to each other.
 5. The turbomachine engine of claim2, wherein the ring gear comprises a first ring gear set that mesheswith the third gear set and a second ring gear set that meshes with thefourth gear set.
 6. The turbomachine engine of claim 1, wherein a gearratio of the gear assembly ranges from 6:1 to 14:1.
 7. The turbomachineengine of claim 1, wherein the gear assembly is a star gearconfiguration in which a planet carrier does not rotate.
 8. Theturbomachine engine of claim 1, wherein the gear assembly is a planetarygear configuration in which the ring gear does not rotate.
 9. Theturbomachine engine of claim 1, wherein the fan assembly is a singlestage of unducted fan blades.
 10. The turbomachine engine of claim 1,wherein the first stage planet gear has a first diameter and the secondstage planet gear has a second diameter, wherein a ratio of the firstdiameter to the second diameter ranges from 1.0 to 2.0.
 11. Theturbomachine engine of claim 2, further comprising two rows of rollerbearings that support, respectively, the third gear set and the fourthgear set of the second stage planet gear.
 12. The turbomachine engine ofclaim 1, wherein there are three planet gear layshafts.
 13. A gearassembly for use with a turbomachine, the gear assembly comprising: asun gear; a plurality of planet gear layshafts that each support a firststage planet gear and a second stage planet gear; and a ring gear, theplanet gear layshaft comprising a first outer portion, a second outerportion, and an intermediate portion between first and second outerportions, wherein the first stage planet gear comprises a first gear setand a second gear set supported at the intermediate portion of theplanet gear layshaft, and the second stage planet gear comprises a thirdgear set supported at the first outer portion of the planet gearlayshaft and a fourth gear set supported at the second outer portion ofthe planet gear layshaft, and the plurality of planet gear layshaftshave a tubular shape.
 14. The gear assembly of claim 13, wherein teethof the first gear set are angularly clocked by a first amount of gearpitch with regards to teeth of the second gear set.
 15. The gearassembly of claim 14, wherein the first amount is between one fourth andone half.
 16. The gear assembly of claim 13, wherein teeth of the thirdgear set are angularly clocked by a second amount of gear pitch withregards to teeth of the fourth gear set.
 17. The gear assembly of claim16, wherein the second amount is between one fourth and one half. 18.The gear assembly of claim 13, further comprising two rows of rollerbearings that support, respectively, the third gear set and the fourthgear set of the second stage planet gear.
 19. The gear assembly of claim18, where the roller bearings comprise ceramic material.
 20. The gearassembly of claims 18, wherein the roller bearings are lubricated underrace.